Hold down and release mechanism for a deployable satellite solar panel

ABSTRACT

The disclosed technology includes systems, methods, and mechanism configurations related to satellite solar panels, including stowing arrangements, deployment sequences, special purpose hinges, hold down and release mechanisms, and associated components for controlled deployment of the satellite solar panels.

FIELD

The disclosed technology relates to satellite solar panel deployment. Inparticular, the disclosed technology includes systems, methods, andmechanism configurations related to satellite solar panels, includinghinges, hold down and release mechanisms, and associated componentsutilized for stowing and controlled deployment of the satellite solarpanels.

BACKGROUND

Nearly all satellites utilize solar panel systems to harvest electricityfor powering on-board systems, charging batteries, etc. The utilizationof solar energy in satellites often requires a tradeoff betweenmaximizing size of the solar panels and minimizing valuable launchvehicle real estate. For example, when a satellite is prepared to belaunched into orbit via a spacecraft, the associated solar panelequipment may be secured in a retracted position to conserve space andto avoid damage to the equipment or the deployer, and to mitigate thechances of orbital debris caused by deployment. Once the satellite is inorbit, the solar panels are unfolded and deployed to an operativeposition.

FIG. 1 depicts a top-view example representation of a prior artsatellite 102 with an “accordion”-style solar panel array 104arrangement in stowed (A), unfolding (B), and deployed (C) positions. Insuch systems, the array 104 solar panels are pivotally coupledend-to-end, folded accordion style for stowage (A), and secured withtie-down/release mechanisms 106, 108 along both edges of the arrayand/or at a central portion of the panels. Once the satellite is inorbit, the restraining mechanisms are released to allow the array 104 tounfold into a substantially planar (or linear) configuration to receivephotons from the sun.

In many prior art systems, the array 104 is attached at a midpoint 110on the body of the satellite 102. In such systems, a half-width panel112 may be utilized to allow for folding the array 104 against one sideof the satellite 102 during stowage (A), but such an arrangement mayleave a portion of unused volume 114 between the folded array 104 andthe body of the satellite 102. Furthermore, the use of a reduced sizepanel 112 (due to the mid-point attachment 110) may reduce the effectivesize (and power harvesting potential) of the unfolded array 104.

One of the drawbacks with accordion-style arrangements and/or attachmentpoints of the conventional system is that the edges of the stowed panel104 typically need to be secured on both respective sides (and/or thecenter) of the satellite body by the tie-down/release mechanisms 106,108, which can create extra bulk. In such prior art systems, the chancesfor a deployment fault can be increased due to the added coordinationneeded for releasing both tie-down/release mechanisms 106, 108. In somesystems, one or more of the tie-down/release mechanisms 106, 108 maybecome detached during deployment, and may create additional orbitaldebris and/or may damage the panel 104 or satellite 102.

A need exists for improved systems and methods to address suchchallenges.

BRIEF SUMMARY

Some or all the above needs may be addressed by certain embodiments andimplementations disclosed herein.

Certain implementations of the disclosed technology may include a systemthat includes a satellite body. The satellite body may define: first,second, third, and fourth major sides; with first, second, third andfourth edge regions disposed between the corresponding first and second,second and third, third and fourth, and fourth and first major sides.The system includes a first solar array having a plurality of panelsattached to one another therebetween by corresponding first-type hingemechanisms. A first panel of the plurality of panels is attached to thesatellite body at the first edge region by one or more second-type hingemechanisms, and at least the first panel of the plurality of panels isreleasably constrained by a hold down and release mechanism (HDRM). Thefirst solar array is configured to be foldable into a stowedconfiguration, and (planar) unrollable into a deployed configuration.

Certain implementations of the disclosed technology may include a methodof planar rolling, into a stowed configuration for stowage against afirst major side of a satellite body, a plurality of solar panels, theplurality of solar panels may be attached to one another therebetween bycorresponding first-type hinge mechanisms. The method includesattaching, by one or more second-type hinge mechanisms, and at or near afirst edge region of the satellite body, a first panel of the pluralityof solar panels, and releasably constraining, by a hold down and releasemechanism (HDRM), the planar rolled plurality of solar panels at an edgeregion adjacent to the first edge region. The (planar) “rolled”plurality of solar panels is configured to be released by the HDRM and(planar) “unrolled” into a deployed linear array configuration having aninnermost panel, a middle panel, and an outermost panel, and theoutermost panel of the deployed linear array is disposed between aninnermost panel and a middle panel of the plurality of panels in thestowed configuration.

Certain implementations of the disclosed technology may include anothermethod of deploying, from a stowed position against a first major sideof a satellite body, a plurality of solar panels having a deployableinnermost, middle, and outermost panel. The deployable outermost panelof the plurality of solar panels is disposed in a first stackedconfiguration between the deployable innermost panel and the deployablemiddle panel during the stowed position. The plurality of solar panelsmay be configured to be rotatably attached to one another therebetweenby corresponding first-type hinge mechanisms, and the deployableinnermost panel of the plurality of solar panels is configured to attachto the satellite body at or near a first edge region by one or moresecond-type hinge mechanisms. A hold down and release mechanism (HDRM)disposed at a second edge region of the satellite body is configured toconstrain and selectively release the plurality of solar panels from thestowed position. The method of deploying can include: releasing from thestowed position, and by the HDRM, a first constrained portion of theplurality of solar panels; rotating the released portion of theplurality of solar panels to a first predetermined angle with respect tothe first major side of the satellite body while maintaining the firststacked configuration; releasing, at the first predetermined angle, andby a first cam and a first stop mechanism, a second constrained end ofthe deployable middle and outermost panel; rotating the released portionof the deployable middle and outermost panel to a second predeterminedangle with respect to the deployable innermost panel while maintaining asecond stacked configuration of the deployable middle panel anddeployable outermost panel; and releasing, at the second predeterminedangle, and by a second cam and a second stop mechanism, a thirdconstrained end of the deployable outermost panel.

Certain implementations of the disclosed technology may include hingeassembly configured to rotatingly join an inboard panel with an outboardpanel of a deployable satellite solar array. The hinge assemblyincludes: a first rotor having a pin portion, a rotor endstop portion,and a first rotor body portion. The first rotor body portion isconfigured to attach to the outboard panel. The hinge assembly includesa first bushing having an inner surface configured to rotatingly engagewith an outer surface of the pin portion of the first rotor. The hingeassembly includes a stator having a first bore, a stator endstopportion, a stator body portion, and a second bore. An inner surface ofthe first bore is configured to engage with an outer surface of thefirst bushing, and the stator endstop portion is configured to restricta relative rotation between the inboard and outboard panel when engagedwith the rotor endstop portion. The stator body portion is configured toattach to the inboard panel. The hinge assembly includes a secondbushing having an outer surface configured to engage with an innersurface of the second bore of the stator. The hinge assembly includessecond rotor having a second rotor pin portion, a spring retentionportion and a second rotor body portion. An outer surface of the secondrotor pin portion is configured to rotatingly engage with an innersurface of the second bushing, and the second rotor body portion isconfigured to attach to the outboard panel. The hinge assembly includesa spring having a coiled portion, a first spring end, and a secondspring end. The coiled portion is configured to surround the springretention portion of the second rotor, the first spring end isconfigured to engage with the stator, and the second spring end isconfigured to engage with the second rotor.

Certain implementations of the disclosed technology may include holddown and release mechanism (HDRM) configured to secure and selectivelyrelease one or more panels of a deployable satellite solar array. TheHDRM can include: a circuit board electrically coupled to an activationcircuit; at least two cylindrical burn resistors electrically coupled tothe circuit board; a burn wire (such as monofilament) in contact withthe at least two cylindrical burn resistors; and two or more restraininghooks configured to retain the burn wire against the at least twocylindrical burn resistors. The activation circuit is configured toselectively direct current through the at least two cylindrical burnresistors to sever the burn wire.

Certain implementations of the disclosed technology may include methodof deploying a satellite solar array. The method can include: securingone or more panels of the satellite solar array against a body of thesatellite with a burn wire; selectively energizing, with an activationcircuit, at least two cylindrical burn resistors in a contact with theburn wire; severing the burn wire by the selectively energizing;releasing by the severing, at least the secured end of the one or morepanels the satellite solar array; and planar unrolling the satellitesolar array into a deployed configuration.

Other implementations, features, and aspects of the disclosed technologyare described in detail herein and are considered a part of the claimeddisclosed technology. Other implementations, features, and aspects canbe understood with reference to the following detailed description,accompanying drawings, and claims.

BRIEF DESCRIPTION OF THE FIGURES

Reference will now be made to the accompanying figures and flowdiagrams, which are not necessarily drawn to scale, and wherein:

FIG. 1 depicts a top-view example representation of a prior artsatellite 102 with an “accordion”-style solar panel array 104arrangement in stowed (A), unfolding (B), and deployed (C) positions.

FIG. 2 is an example illustration of a satellite 200 deployed into orbitrelative to earth 202 and configured to receive solar energy from thesun 204, according to an example implementation of the disclosedtechnology.

FIG. 3A is a top-view illustration of a satellite 200 in a stowedconfiguration, according to an example implementation of the disclosedtechnology.

FIG. 3B is a perspective view illustration of a satellite 200 in astowed configuration, according to an example implementation of thedisclosed technology.

FIG. 4A is an example top-view illustration of the satellite 200 withthe arrays 304, 306 shown in the various stages (a), (b), (c) of anunfolding sequence, according to an example implementation of thedisclosed technology.

FIG. 4B is an example perspective view illustration of the satellite 200with one array 306 shown in the various stages (a), (b), (c) of anunfolding sequence (corresponding to the stages shown in FIG. 4A),according to an example implementation of the disclosed technology.

FIG. 5A is a top-front perspective view depiction of a deployedsatellite 200, according to an example implementation of the disclosedtechnology.

FIG. 5B is top-rear view perspective depiction of a deployed satellite200, according to an example implementation of the disclosed technology.

FIG. 6A depicts the satellite 200 in an unfolding stage (as depicted inthe corresponding initial unfolding stage (a) of FIGS. 4A and 4B) inwhich the cam 604 remains constrained by the stop 606.

FIG. 6B depicts the panels 608, 610, 612 in a stowed configuration,according to an example implementation of the disclosed technology.

FIG. 6C depicts the panels 608, 610, 612 in a stage of rotation (similarthe stage of rotation shown in FIG. 6A) before the first predeterminedangle of rotation has been exceeded.

FIG. 6D depicts the panels 608, 610, 612 in a mid-deployment stage inwhich the innermost panel 608 has rotated past the first predeterminedrotation angle, allowing the offset cam 604 to clear the stop 606,thereby allowing the middle panel 610 and the outermost panel 612 toswing clear of the innermost panel 608 and the body of the satellite200.

FIG. 7A depicts the satellite 200 in an unfolding stage (as depicted inthe corresponding second unfolding stage (b) of FIGS. 4A and 4B) inwhich the cam 704 remains constrained by the stop 706.

FIG. 7B depicts the panels 608, 610, 612 during a first stage ofunfolding (such as described above and as depicted in FIG. 6A or FIG.6C), according to an example implementation of the disclosed technology.

FIG. 7C depicts the panels 608, 610, 612 near a second stage ofrotation, i.e. right before the second predetermined angle of rotationhas been exceeded.

FIG. 7D depicts the panels 608, 610, 612 in a final-deployment stage inwhich the innermost panel 608 has rotated past the first predeterminedrotation angle, and where the middle panel 610 and outermost panel 612have rotated past the second predetermined rotation angle, allowing thecam 704 to clear the stop 406, thereby allowing the outermost panel 612to swing clear of the innermost panel 608 and the body of the satellite200.

FIG. 8A is a side face view illustration of a planar rolled-up solararray 306, according to an example implementation of the disclosedtechnology.

FIG. 8B is a side edge view illustration of the rolled-up array 306 in astowed configuration.

FIG. 9A is an exploded-view illustration of a hinge assembly 308,according to an example implementation of the disclosed technology.

FIG. 9B is a cross-sectional side-view illustration of the hingeassembly 308.

FIG. 10 illustrates a partial view of a solar pane array 306 and anexample mirrored pair positioning and orientation of the hingeassemblies 308 with respect to the innermost panel 608 and the middlepanel 610, according to an example implementation of the disclosedtechnology.

FIG. 11A is a side-view perspective illustration of a hold down andrelease mechanism (HDRM) 310 (as also depicted in FIGS. 3A, 3B, 8A, and8B) according to an example implementation of the disclosed technology.

FIG. 11B is a cross sectional side-view illustration of the HDRM 310,with like components as depicted in FIG. 11A.

FIG. 12 is a cross sectional side-view illustration of another exampleimplementation of a hold down and release mechanism (HDRM2) 1200.

FIG. 13 depicts an example activation circuit 1300 that may be utilizedto selectively control a current 1301 through the burn resistors 1102 tosever the monofilament 1106 for controllable deployment of the solarpanels.

FIG. 14 is a flow diagram of a method 1400, according to an exampleimplementation of the disclosed technology.

FIG. 15 is a flow diagram of a method 1500, according to an exampleimplementation of the disclosed technology.

FIG. 16 is a flow diagram of a method 1600, according to an exampleimplementation of the disclosed technology.

DETAILED DESCRIPTION

Some implementations of the disclosed technology will be described morefully hereinafter with reference to the accompanying drawings. Thisdisclosed technology may however, be embodied in many different formsand should not be construed as limited to the implementations set forthherein. Some of the components illustrated in the accompanying figuresare shown for illustration purposes only, and may or may not be drawn toscale. In the following detailed description, numerous specific detailsare set forth by way of examples to provide a thorough understanding ofthe relevant teachings. However, it should be apparent to those skilledin the art that the present teachings may be practiced without suchdetails. In other instances, well known methods, procedures, components,and/or circuitry have been described at a relatively high-level, withoutdetail, to avoid unnecessarily obscuring aspects of the presentteachings.

As used herein, the terms planar “rollable,” “rolling,” “rolled,” and/or“unrollable,” “unrolling,” “unrolled” (or equivalents) may include anystructure or associated action in which a structure is sequentiallyfolded and/or unfolded in sections in a manner similar to rolling.

FIG. 2 is an example illustration (not to scale) depicting a setting ofthe disclosed technology in which a satellite 200 (having unfolded solarpanels) is deployed into orbit relative to earth 202 and configured toreceive solar energy from the sun 204. According to certain exampleimplementations of the disclosed technology, the satellite 200 may be aCubeSat deployed into low earth orbit. As defined herein, and accordingto an example implementation of the disclosed technology, a CubeSat maybe generally defined as a class of nano-satellites, with dimensionsapproximately 10 cm×10 cm×10 cm for each 1 unit (referred to in theindustry as 1U, being the basic building block of CubeSats which can bearranged in 1U, 2U, 3U, 6U, or other formats), and having approximately1.5 kg of mass per 1U. For example, a 3U CubeSat would be approximately10 cm×10 cm×30 cm and 4.5 kg, being an arrangement of three 1U CubeSats.While a 3U CubeSat is depicted and discussed herein, the disclosedtechnology is not limited to 3U CubeSats, and may be applied to a widevariety of other types of satellites in other associated orbits,including other types of CubeSats.

As will be discussed herein, various aspects of the disclosed technologymay provide certain technical features and benefits associated withdeploying a satellite 200 into orbit, including but not limited to:terrestrial preparation and testing, folding and securing the solararrays to the body of the satellite 200, reducing volume of the stowedsatellite 200, providing standby power via exposed solar cells whenstowed, reliably releasing and deploying the folded solar arrays when inorbit, etc.

Stowing and Unfolding of Solar Panels

FIG. 3A and FIG. 3B depict respective top-view and perspective viewillustrations of a satellite 200 in a stowed configuration, according toan example implementation of the disclosed technology. As shown in thetop-view FIG. 3A, a relative orientation (for descriptive purposes) ofthe first, second, third, and fourth major sides of the satellite 200may be denoted respectfully as 1MS, 2MS, 3MS, and 4MS. For example, FIG.3B is depicted with the forth major side 4MS closest to the viewer.Also, as shown in the top-view FIG. 3A are first, second, third andfourth edge regions denoted respectfully as 1E, 2E, 3E, and 4E. Theseedge regions (as depicted, for example, by the dashed circle around thefirst edge region 1E) are disposed between the corresponding first andsecond, second and third, third and fourth, and fourth and first majorsides. For example, FIG. 3B is depicted with the fourth edge (4E)between the fourth and first major sides. In certain exampleimplementations, the edge regions may include additional features,shapes, cutouts, etc. In certain example implementations, the “edgeregions” may comprise regions at or near the corners of the satellite200. For example, the edge region may extend from a center of a majorside face to a corresponding edge (i.e., 0-50% of the width of the majorside face). In certain example implementations, the edge region mayinclude 0-25 percent of the width of the satellite 200. In certainexample implementations, the edge region may include 0-10 percent of thewidth of the satellite 200. In certain example implementations of thedisclosed technology, the edge region may be defined such that thepanels attached within the edge region substantially cover a major sideof the satellite 200.

In this example configuration shown in FIGS. 3A and 3B, the individualsolar panels 302 of the associated arrays 304, 306 are shown folded-upin a planar “rolled” compact arrangement against respective body sidesthe satellite 200 (as will be discussed further below with respect toFIGS. 4A-7D). According to an example implementation of the disclosedtechnology, the panels 302 may be joined with hinge assemblies 308 (forexample, in or near opposite edge regions 1E, 3E), and secured with oneor more hold down release mechanisms (HDRM) 310 (for example, in or nearthe corresponding adjacent edge regions 4E and 2E). In certain exampleimplementations, the hinge assemblies 308 that join the arrays 304 306to the body of the satellite 200 may be configured to include and/or beattached to an extension piece 314, for example, to accommodate thethickness of the outer panels when in the stowed configuration. Anexample implementation of the extension piece 314 is also shown in FIG.6B.

The general configuration of the panels 302, hinge assemblies 308 and/orthe HDRMs 310, as depicted in FIG. 3A and FIG. 3B is shown here forillustration purposes only, and may not be to scale. Exampleimplementations with additional details, including positioning, size,shape, and operation of these components, will be discussed below withrespect to the subsequent figures.

In accordance with certain example implementations of the disclosedtechnology, one or more separate fixed solar panels 312 may be attachedto one or more respective body sides of the satellite 200, and adjacentto the folded arrays 304, 306. According to an example implementation ofthe disclosed technology, the fixed solar panels 312 may enable thesatellite 200 to collect solar power when stowed and/or deployedproviding at least some power in the event of a deployment failure.

FIG. 4A and FIG. 4B depict respective example top-view and 3D-viewillustrations of the satellite 200 with the arrays 304, 306 shown in thevarious stages (a), (b), (c) of an unfolding sequence, according to anexample implementation of the disclosed technology. For simplicity, onlyone of the arrays 306 is depicted in FIG. 4B. As illustrated, andaccording to an example implementation of the disclosed technology, thearrays 304, 306 may be joined to the body of the satellite 200 at ornear respective opposite corner edges 402, 404 of the satellite, incontrast with prior art systems in which the arrays are attached at amidpoint (such as midpoint 110 as discussed above with respect to FIG.1). Furthermore, in contrast to prior art systems that utilize an“accordion” folding arrangement (as discussed with respect to FIG. 1),certain example implementations of the disclosed technology provide aplanar “rolled-up wing” structure in which the individual planar panels302 of the arrays 304, 306 each unfold/unroll in a unified direction(such as counter clockwise when viewed from the top, as illustrated inFIG. 4A). According to an example implementation of the disclosedtechnology, the planar “rolled-up wing” structure and unified unfoldingdirection enables the respective arrays 304, 306 to be attached at ornear corner edges 402, 404 of the satellite by hinges 308, and securedwith a reduced number of hold down release mechanisms (HDRMs) 310, asshown in FIGS. 3A and 3B, without requiring such HDRMs 310 to alsosecure the arrays 304, 306 at the corner edges 402, 404. As will besubsequently explained below, certain hinge, cam, and stop features,etc., may be utilized to simplify the stowing and subsequent unfoldingof the arrays 304, 306.

FIG. 5A is a top-front view depiction of a deployed satellite 200,according to an example implementation of the disclosed technology.According to certain example implementations, the satellite 200 may bedeployed in an orientation such that the front sides of the arrays 304,306 having the most solar cells face the sun 204 (not to scale) formaximum solar power harvesting potential.

FIG. 5B is top-rear view depiction of a deployed satellite 200,according to an example implementation of the disclosed technology, inwhich at least a portion of the back sides of the arrays 304, 306 arepopulated with solar cells 502. According to an example implementationof the disclosed technology, this arrangement (of solar cells 502 on theback sides) may allow for a certain amount of solar power harvestingeven when the front side of the arrays 304, 306 of satellite 200 are notfacing sun 204. In such instances, for example, the satellite 200 maystill operate as intended. In certain example implementations, thesatellite 200 may operate with reduced functionality. In certain exampleimplementations, the satellite 200 may continue to harvest solar powervia the solar cells 502 on the back sides of the arrays 304, 306 and/orone or more separate fixed solar panels 312. In certain implementation,the partially-populated back sides of the arrays 304, 306 and/or one ormore separate fixed solar panels 312 may provide enough power (whenfully or partially facing the sun) to charge batteries, powerelectronics, power orientation correction propulsion systems, operatecommunication radios, etc. Thus, certain implementations of thedisclosed technology may enable harvesting solar power, even in caseswhere the satellite 200 is deployed in an undesirable orientation withrespect to the sun 204.

Unfolding Deployment Control Mechanisms

FIGS. 6A-6D depict the operation of an inner panelunfolding/retention/sequencing first stage mechanism 602, according tocertain example implementations of the disclosed technology. Thesatellite 200 depicted in FIG. 6A (left) is shown for simplicity withonly one array 306 unfolding, however, certain implementation mayinclude another foldable/unfoldable array (such as array 304 as shown inFIG. 4A and FIG. 5A). In certain example implementations, the firststage mechanism 602 includes an offset cam 604 and a stop 606 that areconfigured to engage during rotation to prevent the unfolding of themiddle panel 610 and the outermost panel 612 until the innermost panel608 has rotated enough beyond a first predetermined angle, i.e., toprovide sufficient clearance for the rotation of the middle panel 610and the outermost panel 612 without scraping, disturbing, or otherwisetouching the body of the satellite 200. For example, the stop 606 may beconfigured with appropriate geometries and the cam 604 may be configuredwith the appropriate axial offset such that the cam 604 remainsconstrained by the stop 606 until the array 306 rotates greater thanabout 90 degrees from the stowed position (such as the stowed positiondepicted in FIG. 6B). In certain example implementations, the cam 604may be integrated with the panel hinge, as will be explained below.

FIG. 6A depicts the satellite 200 in an unfolding stage (as depicted inthe corresponding initial unfolding stage (a) of FIGS. 4A and 4B) inwhich the cam 604 remains constrained by the stop 606. In certainexample implementations, the stop 606 may be attached to the body of thesatellite 200 and may be configured to engage with the cam 604 duringthe initial unfolding stage. According to an example implementation ofthe disclosed technology, a release mechanism may be utilized toinitiate the unfolding process, and a spring integrated with the hingeassembly may provide the force necessary to unfold the panels, as willbe further explained below with reference to the hinge assembly.

In certain example implementations, the stop 606 may have a taperedshape, as depicted. In other example implementations, the stop 606 mayinclude other shapes and features, including but not limited to notches,surface treatments, etc. For example, in certain implementations, thestop 606 may be configured with a smooth sliding surface to enable easydisengagement from the cam 604 once the innermost panel 608 has unfoldedpast a first predetermined angle.

FIG. 6B depicts the panels 608, 610, 612 in a stowed configuration,according to an example implementation of the disclosed technology.According to an example implementation, the radial axis 614 of the cam604 may be offset from the corresponding radial axis 616 of the hinge.In this configuration, the cam 604 may engage with the stop 606 tosecure the middle panel 610 and the outermost panel 612 against oradjacent to the innermost panel 608. In accordance with certain exampleimplementations of the disclosed technology, the cam 604 may be attachedto a hinge assembly and disposed between the middle panel 610 and theoutermost panel 612. In certain example implementations, the hingeassemblies 308 that join the innermost panel 608 to the body of thesatellite 200 may be configured to include and/or be attached to anextension piece 314, for example, to accommodate the thickness of thepanels 608, 610, 612 when in the stowed configuration.

FIG. 6C depicts the panels 608, 610, 612 in a stage of rotation (similarto the stage of rotation shown in FIG. 6A) before the firstpredetermined angle of rotation has been exceeded. In this position, thecam 604 is still engaged with the stop 606, thereby holding the middlepanel 610 and the outermost panel 612 adjacent to the innermost panel608.

FIG. 6D depicts the panels 608, 610, 612 in a mid-deployment stage inwhich the innermost panel 608 has rotated past the first predeterminedrotation angle, allowing the offset cam 604 to clear the stop 606,thereby allowing the middle panel 610 and the outermost panel 612 toswing clear of the innermost panel 608 and the body of the satellite200.

FIGS. 7A-7D depict the operation of an unfolding/retention second stagemechanism 702, according to an example implementation of the disclosedtechnology. The satellite 200 depicted in FIG. 7A (left) is shown forsimplicity with only one array 306 unfolding, however, certainimplementations may include another foldable/unfoldable array (such asarray 304 as shown in FIG. 4A and FIG. 5A). In certain exampleimplementations, the second stage mechanism 702 includes a cam 704 and astop 706 that are configured to engage during rotation to prevent theunfolding of the outermost panel 612 until the middle panel 610 hasrotated enough beyond a second predetermined angle, i.e., to providesufficient clearance for the rotation of the outermost panel 612 withoutscraping, disturbing, or otherwise touching the body of the satellite200. For example, the stop 706 and cam 704 may be configured withappropriate geometries such that the cam 704 remains constrained by thestop 706 until the engaged middle panel 610 and outermost panel 612rotate together to about 180 degrees away from the face of the innermostpanel 608.

FIG. 7A depicts the satellite 200 in an unfolding stage (as depicted inthe corresponding second unfolding stage (b) of FIGS. 4A and 4B) inwhich the cam 704 remains constrained by the stop 706. In certainexample implementations, the stop 706 may be attached to the innermostpanel 608 and may be configured to engage with the cam 704 during thefirst stage of unfolding, and also during the initial portion of secondstage of unfolding. In certain example implementations, the cam 704 maybe attached to the outermost panel 612.

In certain example implementations, the cam 704 may be cylindrical, asdepicted. In other example implementations, the cam 704 may includeother shapes and features, including but not limited to notches, surfacetreatments, etc. For example, in certain implementations, the cam 704may be configured with a smooth sliding surface to enable easydisengagement from the stop 706 once the middle panel 610 has unfoldedpast the second predetermined angle.

FIG. 7B depicts the panels 608, 610, 612 during a first stage ofunfolding (such as described above and as depicted in FIG. 6A or FIG.6C), according to an example implementation of the disclosed technology.In this configuration, the cam 704 may engage with the stop 706 tosecure the middle panel 610 against (or adjacent to) the outermost panel612.

FIG. 7C depicts the panels 608, 610, 612 near a second stage ofrotation, i.e. right before the second predetermined angle of rotationhas been exceeded. In this position, the cam 704 is still engaged withthe stop 706, thereby securing the outermost panel 612 foldedtowards/against the middle panel 610.

FIG. 7D depicts the panels 608, 610, 612 in a final-deployment stage inwhich the innermost panel 608 has rotated past the first predeterminedrotation angle, and where the middle panel 610 and outermost panel 612have rotated past the second predetermined rotation angle, allowing thecam 704 to clear the stop 706, thereby allowing the outermost panel 612to swing clear of the innermost panel 608 and the body of the satellite200.

As depicted in FIG. 8A and FIG. 8B, an array 306 (shown folded in FIG.8B) may be held to the body of the satellite 200 by hinges 308 along oneedge, while hold down release mechanisms (HDRMs) 310 (see FIGS. 3A and3B) can be placed along the opposite edge of a folded array 306 so thatit may be effectively restrained when stowed. In addition, and accordingto certain example implementations, spacers 802 may be placed betweenthe satellite 200, and each of the panels 608, 610, 612 of the array306, providing more protection from vibration and shock. In certainexample implementations, the entire stack of panels 608, 610, 612 may besecured firmly against the mounting body of the satellite, which mayprovide certain protections for the fragile solar cells. Since thehinges 308 and HDRM 310 points are disposed along the edges of thepanels, there is more available surface area on the panel forcomponents, including solar cells. In certain implementations, the abovefeatures may result in a resilient package when stowed, which may allowuse of thinner panels that can be lighter and can be packed moredensely. Another beneficial aspect of stowing the panels 608, 610, 612of the array 306 in the planar rolled configuration is that sucharrangement allows the innermost panel 608 to surround and constrain theother panels 610 612 against the satellite 200 when stowed so that thefirst hinge 308 secures the array 306 against one edge, while the HDRMsecures the array 306 against the adjacent edge. In contrast,accordion-type arrangements may require HDRMs along both edges (inaddition to hinges). Thus, certain aspects of the disclosed technologyprovide deployable satellite solar panels with reduced components,reduced associated weight, and reduced bulk.

Hinge Assembly

FIG. 9A is an exploded-view illustration of a hinge assembly 308,according to an example implementation of the disclosed technology. FIG.9B is a cross-sectional view of the hinge assembly 308. Certain exampleimplementations of the hinge assembly 308 may be utilized to join aninboard panel (such as the middle panel 610 shown in FIG. 6D) with anoutboard panel (such as the outermost panel 612 shown in FIG. 6D). Incertain example implementations, a similarly configured hinge assembly308 may be utilized to join the innermost panel of the array (such aspanel 608 as shown in FIG. 6D) to the body of the satellite 200.

In accordance with certain example implementations of the disclosedtechnology, the hinge assembly 308 may include one or more of thefollowing components: a first rotor 902 with an integrated end stop 903,a bushing 904, a stator 906 with an integrated end stop 907, a flangedbushing 908, a spring 910 having a first leg section 912 and a secondleg section 914, and a second rotor 916. As shown in FIG. 9B, theassociated components 902-916 of the hinge assembly 308 may be attachedby fasteners 918 to respective portions of satellite body, inboardpanels and/or outboard panels (for example, panel 610 and/or panel 612)of the arrays.

According to an example implementation of the disclosed technology, thefirst rotor 902 may act as a mounting structure to the outboard panel,and may also act as a rotating pin with the mechanical end stop 903 forconstraining the hinge rotation. The end stop 903, for example may set adeployed position of the panel without relying on additional hardware tobe mounted to the panels. The bushing 904, for example, may provide asliding surface for the rotation of the rotor 902. The stator 906, forexample, may act as a mounting structure for the inboard panels (forexample, panel 610 and/or panel 608), and the integrated end stop 907 ofthe stator 906 may engage with an end stop 903 portion of the firstrotor 902, for example, to limit the rotational range of the hinge 308.The flanged bushing 908, for example, may provide a sliding surface forthe rotation of the second rotor 916 which may constrain the hingeaxially. The second rotor 916, for example, may act as a rotating pinand mounting structure for attachment to the outboard panels (forexample, panel 610 and/or panel 612) of the arrays. The spring 910, forexample may provide stored energy and actuation force for the deploymentof the associated panels.

In certain example implementations, the hinges 308 may be designed withtwo rotational range variants: 135 degrees; and 180 degrees. Forexample, the hinges 308 configured with the 135-degree rotational rangemay be attached to the satellite body 200 and the inner edge of theinnermost panel 608, while the 180 degree variants may be utilized tojoin the middle panel 610 with the innermost panel 608 and the outermostpanel 612.

In accordance with certain example implementations of the disclosedtechnology, the integrated spring 910 in each hinge assembly 308 may beutilized to supply the opening force for deployment of the associatedsolar panels. In certain example implementations, first leg section 912of the spring 910 may be restrained by the stator 906, while the secondleg section 914 may be restrained by the second rotor 916.

In certain example implementations of the disclosed technology, theintegrated end stop 903 of the first rotor 902 and the integrated stop907 of the stator 906 may be configured to fix the final deployedposition of the array. The spring 910, for example, may be configured toprovide not only the opening force for the associated panels, but mayalso provide the adequate force to retain the panels in the deployed(open) position when the respective end stop 903 of the first rotor 902is engaged with end stop 907 the stator 906.

According to certain example implementations of the disclosedtechnology, one or more of the components of the hinge assembly 308 maybe sized as needed. For example, the length of the stator 906 may beconfigured differently depending on the intended placement of the hingeassembly 308. In other example implementations, each of the hingeassemblies 308 may have a substantially similar geometry. According toan example implementation of the disclosed technology, each hinge mayfit within a defined volume along the side of the panel withoutintruding onto a volume defined by the associated panel.

Certain example implementations of the disclosed technology utilizelow-profile screws as fasteners 918 to attach to the panels, which mayhelp avoid certain inaccuracies in panel mounting hole placement, andmay also help avoid certain frictions in the operation of the hinge 308.

According to an example implementation of the disclosed technology, thehinge assemblies 308 may be contained within the area along the edge ofthe respective panels such that they do not impinge on the center areaof the panel. According to an example implementation of the disclosedtechnology, the functional elements 902-918 of the installed hingeassemblies 308 may be configured to occupy a volume at the edge of thepanels to provide maximal surface area for components such as solararrays. In certain example implementations, the mechanical design of thepanel may be decoupled from the electrical design and layout of thepanel solar cells.

FIG. 10 illustrates a partial view of a solar pane array 306 and anexample mirrored pair positioning and orientation of the hingeassemblies 308 with respect to the innermost panel 608 and the middlepanel 610, according to an example implementation of the disclosedtechnology.

Hold Down Release Mechanism (HDRM)

FIG. 11A is a 3D illustration of a hold down and release mechanism(HDRM) 310 (as also depicted in FIGS. 3A, 3B, 8A, and 8B) according toan example implementation of the disclosed technology. FIG. 11B is across sectional side-view illustration of the HDRM 310, with likecomponents as depicted in FIG. 11A.

According to an example implementation of the disclosed technology, theHDRM 310 may include one or more cylindrical burn resistors 1102 thatare configured to heat up when current passes through them, therebysevering the burn wire 1106 (herein also referred to as monofilament1106) and allowing the panels 608, 610, 612 to unfold and deploy via theforce of the springs (such as spring 910 as shown in FIGS. 9A and 9B. Incertain example implementations, the cylindrical burn resistors 1102 maybe mounted to a circuit board 1110, which in certain implementations,may also act as a thermal isolation panel to at least partiallyconstrain the heat generated when current flows through the burnresistors 1102. In one example implementation of the disclosedtechnology, the burn resistors 1102 may be mounted to a surface ofcircuit board 1110. In another example implementation, the burnresistors 1102 may be “pigtail” style and mounted to the circuit board1110 via leads, through holes, etc.

According to an example implementation of the disclosed technology,restraining hooks 1104 may be mounted to the circuit board 1110 anddisposed on either side of the cylindrical burn resistors 1102 to act asrestraining guides for a monofilament 1106, and to increase theeffective contact surface area between the monofilament 1106 and burnresistors 1102. According to an example implementation of the disclosedtechnology, the monofilament 1106 may be threaded through therestraining hooks 1104, tightened, and secured at both ends by hold-downscrews 1108 such that the monofilament 1106 makes intimate contact withone or more of the cylindrical burn resistors 1102.

In accordance with certain example implementations of the disclosedtechnology, one of more additional monofilament segments 1107 (asdepicted with dashed lines in FIG. 11A) may be utilized to furthersecure the panels against the satellite body 200. The additionalmonofilament segments 1107, for example, may provide additional tie-downstrength for stowage of the associated panels without requiringadditional burn resistors 1102. In certain example implementations, theadditional monofilament segments 1107 may be secured with common holddown screws 1108. In other certain example implementations, theadditional monofilament segments 1107 may be secured with separate holddown screws (not shown). In some example implementations, the additionalmonofilament segments 1107 may be looped and tied to itself (in asimilar fashion to the monofilament 1208 as discussed below and shown inFIG. 12).

In certain example implementations, the circuit board 1110 may besecured to a top surface of the outer panel 608 and co-located with aclearance (for example, a recess in one or more of the panels 608, 610,612) such that the tightened monofilament 1106 does not touch or rub onedges of the panels 608, 610, 612.

In accordance with certain example implementations of the disclosedtechnology, one or more washers may be placed between the hold-downscrews 1108 and the monofilament 1106 to allow rotation of the hold-downscrews 1108 during tightening while avoiding damage to the monofilament1106. In certain example implementations, the washers may be made of acompressible material (such as rubber or plastic) to provide enhancedgripping friction, and to reduce slipping of the monofilament 1106 whensecured by the hold-down screws 1108. In certain exampleimplementations, the washers may include texturing on one side, and asmooth surface on the other side, so that the textured side may securemonofilament 1106 against the satellite body 200, while the smooth sideprovides slippage of the hold-down screws 1108 during tightening. Incertain example implementations, the washer and/or the hold-down screws1108 may include certain locking features that may help prevent thehold-down screws 1108 from becoming loose during transport, vibration,etc. In certain example implementations, the hold-down screws 1108 maybe secured with an adhesive and/or other thread treatments.

In certain example implementations, the restraining hooks 1104 may bemade from a metal material. In certain example implementations, therestraining hooks 1104 may include wire bent in a u-shape, fed intoplated through holes in the circuit board 1110, and soldered to thecircuit board 1110. In certain example implementations, the restraininghooks 1104 may have ends that are bent, flattened, and/or furthersecured to minimize the possibility of coming disengaged duringtransport. Certain example implementations of the restraining hooks 1104may be of sufficient diameter and/or surface smoothness such that themicrofilament 1106 is not frayed or otherwise severed, for example, byvibrations during transport.

According to an example implementation of the disclosed technology, thesecured monofilament 1106 may be used to restrain the stowed panels 608,610, 612. For example, the outer stowed panel 608 may be held againstthe other folded (planar rolled-up) panels 612, 610 (via spacers) andagainst a portion of the satellite body 200 (or against anotherappropriate stop portion) by the tightened and secured microfilament1106, for example to secure the folded array (such as array 306) againstthe satellite 200 for stowage and pre-deployment.

In certain example implementations, the panels 608, 610, 612 may befolded and secured with the HDRM 310 such that a stow detect switch 1116is engaged. In certain example implementations, the stow detect switch1116 may be utilized to provide a signal indicative of whether thepanels are in a stowed position. In certain example implementations ofthe disclosed technology, the stow detect switch 1116 may be positioneddirectly adjacent to or near the HDRM 310 (for example, at the edge ofthe outer panel 608 as shown in FIG. 11A) so that any switch engagementerrors due to panel flexing, etc. may be minimized.

In accordance with certain example implementations of the disclosedtechnology, to release the folded array, an appropriate current may bedirected through one or more of the burn resistors 1102. The associatedresistive heating may be sufficient to melt through the monofilament1106, thereby releasing the stowed panels and allowing the solar arrayto planar unroll (in the controlled fashion as discussed above withrespect to FIGS. 6A-6D and FIGS. 7A-7D) and open for deployment (asshown in FIG. 5A, for example).

In certain example implementations, the HDRM 310 may include two or moreburn resistors 1102 to provide redundancy. In accordance with certainexample implementations of the disclosed technology, the resistors 1102may be wired in parallel (rather than in series) so that at least one ofthe burn resistors 1102 can conduct current, even if the other onebecomes damaged or open circuited.

In accordance with certain example implementations, the monofilament1106 (as also referred to herein as a “burn wire”) may be commerciallyavailable fishing line, and may be selected based on tensile strength,melting temperature, etc. In one example implementation of the disclosedtechnology, the monofilament 1106 may be a 0.12 mm diameter Dyneema®(ultra-high-molecular-weight polyethylene), or similar material whichsufficiently weakens at temperatures at approximately 60° C. and aboveto allow deployment of the mechanism. In certain exampleimplementations, the monofilament 1106 may be selected, as appropriate,to have a diameter that may range from about 0.05 mm to about 0.3 mm.For example, when additional segments 1107 of the monofilament 1106 areutilized in the HDRM 310, the diameter of the monofilament 1106 may bereduced. In certain example implementations, the number of additionalsegments 1107 of the monofilament 1106 and/or the thickness of themonofilament 1106 may be selected to provide appropriate retentionstrength while reliably melting during deployment. In certain exampleimplementations, the diameter of the monofilament 1106 (and/oradditional segments 1107) may be selected based on the availableprojected temperature rise of the burn resistors 1102 during deployment.In certain example implementations, the HDRM 310 may include additionalthermal shielding (not shown) around the burn resistors 1102 to reduceheat dissipation and increase the available temperature rise associatedwith the burn resistors 1102 during deployment.

FIG. 12 is a cross sectional side-view illustration of another exampleimplementation of a hold down and release mechanism (HDRM2) 1200, inwhich an outer stowed solar panel 1202 may be secured, for example,against the satellite body 1204 and/or against an inner panel 1206 thatmay be attached to the satellite body 1204. In certain exampleimplementations of the disclosed technology, the inner panel 1206 may beconfigured to include solar cells (such as the fixed solar panels 312shown in FIG. 5B.

According to an example implementation of the disclosed technology, theHDRM2 1200 may include one or more cylindrical burn resistors 1102, aspreviously discussed, that are configured to heat up when current passesthrough them, thereby severing a burn wire 1208 (such as themonofilament 1106 shown in FIG. 11A), and allowing the outer stowedpanel 1202 to unfold and deploy.

According to an example implementation of the disclosed technology,restraining pins 1210 may be disposed on either side of the cylindricalburn resistors 1102 to act as restraining guides for the burn wire 1208,and to increase the effective contact surface area between the burn wire1208 and burn resistors 1102. According to an example implementation ofthe disclosed technology, the burn wire 1208 may be installed,tightened, and secured, for example, by a knot, adhesive, and/ordeformable collar such that the burn wire 1208 makes extensive contactwith one or more of the cylindrical burn resistors 1102 and restrainsthe outer stowed solar panel 1202, for example, against one or moreholdoffs or spacers 1212 between the panels 1202 1206.

FIG. 13 depicts an example activation circuit 1300 that may be utilizedto selectively control a current 1301 through the burn resistors 1102 tosever the monofilament 1106 (and/or additional/optional monofilamentsegment 1107) for controllable deployment of the solar panels. Forexample, a power source 1302 (such as a battery) may be connected to theburn resistors 1102 via a controllable switching device 1304 (such as arelay, solid state switching device, etc.). The activation circuit 1300may be completed via a common ground 1306. According to an exampleimplementation of the disclosed technology, the controllable switchingdevice 1304 may receive an actuation signal 1308 by an on-boardcontroller 1310 such that when selectively activated, current may flowthrough the burn resistor(s) 1102.

According to certain example implementations of the disclosedtechnology, the burn resistors 1102 may be selected based on specificheat capacity, mass, wattage, and resistance so that when the switchingdevice 1304 is actuated to allow current 1301 to flow, the temperatureof the burn resistors 1102 will rise above the melting temperature ofthe monofilament 1106 (˜110° C.) in a short period (i.e., less thanabout 10 seconds). According to an example implementation of thedisclosed technology, the melting temperature of the monofilament 1106may be on the order of 110° C. or above, however there is a temperaturepoint below the melting temperature (experimentally determined betweenapproximately 60° C. and 70° C.) at which certain monofilament 1106 maysignificantly weaken, to the point that it snaps under the force of thehinge spring. For each implementation of the HDRM, trial experimentswith various values of the burn resistors 1102 have been tested todetermine values that will sever the monofilament after approximately 5s. It has been experimentally determined that resistance value ofapproximately 12 ohms to approximately 20 ohms (in 0204 package) for theburn resistors 1102 provides the desired melting of the monofilament1106 for excitation voltage across the burn resistors 1102 ranging fromapproximately 5 volts to approximately 12 volts.

Energy and temperature are related by the heat capacity equation:

${{\Delta \; T} = \frac{\Delta \; Q}{m\mspace{14mu} C}},$

where ΔT is the temperature rise, ΔQ is the net added energy, m is themass of the material and C is the specific heat capacity of the materialper unit mass. As an illustration, consider a single 5 ohm, 10 watt, 2gram, ceramic wire wound resistor (with ceramic having a specific heatof about 1 J/(gK)) and subjected to a 5 volt supply across the resistorfor 10 seconds. In this case, and ignoring losses to the environment, 1amp of current will flow, with an integrated total of 10 joules (ΔQ)applied over the actuation period. The approximate rise in temperatureof the resistor for this case would be about 5° C. While the ratedwattage of this example resistor may be adequate, the associatedtemperature rise of the resistor would obviously not be sufficient tomelt a nylon-based monofilament.

As another illustration, suppose the (2 gram) resistor above is replacedwith a smaller 40 milligram resistor (with all other specs as before).In this example, the rise in temperature could increase by a factor of50, or to about 250° C. However, as may be appreciated by those havingskill in the art, the wattage rating on such a small resistor may not besufficient. According to certain example implementations of thedisclosed technology, the specifications of the burn resistors 1102(i.e., materials, mass, wattage, resistance, etc.) and the voltage andcurrent capacity of the power source 1302 may be selected such that thetemperature of the burn resistors 1102 may be elevated above the meltingpoint of the monofilament 1106 without prematurely destroying the burnresistors 1102 before the monofilament 1106 is severed.

According to certain example implementations of the disclosedtechnology, the burn resistors 1102 may be 15 ohm thin-film MLEFresistor, rated for 250 mW, in a 0204 package. Although target meltingtimes vary with the physical implementation of the HDRM and ambienttemperatures, in an example implementation 8V supply voltage providessufficient heat to melt the monofilament in approximately 5 seconds atroom temperature.

FIG. 14 is flow diagram of a method 1400, according to an exampleimplementation of the disclosed technology. In block 1402, the method1400 includes planar rolling, into a stowed configuration against afirst major side of a satellite body, a plurality of solar panels, theplurality of solar panels are attached to one another therebetween bycorresponding first-type hinge mechanisms. In block 1404, the method1400 includes attaching, by one or more second-type hinge mechanisms,and at a first edge region of the satellite body, a first panel of theplurality of solar panels. In block 1406, the method 1400 includesreleasably constraining, by a hold down and release mechanism (HDRM),the planar rolled plurality of solar panels at an edge region adjacentto the first edge region.

In certain example implementations, the first-type hinge mechanism maybe characterized by a rotation angle range of about 0 when stowed toabout 180 degrees when deployed.

In certain example implementations, the second-type hinge mechanism maybe characterized by a rotation angle range of about 0 when stowed toabout 135 degrees when deployed.

In certain example implementations, the HDRM may be disposed at a fourthedge region of the satellite body, and wherein a first combinedrotatable edge of the plurality of solar panels in the stowed positionis disposed opposite the second-type hinge mechanism.

In certain example implementations, the second-type hinge mechanism mayconstrain the deployed linear array by a hinge spring and a hinge stop.

FIG. 15 is flow diagram of a method 1500, according to an exampleimplementation of the disclosed technology. In block 1502, the method1500 includes releasing from a stowed position, and by a hold down andrelease mechanism (HDRM), a first constrained portion of a plurality ofsolar panels. In block 1504, the method 1500 includes rotating thereleased portion of the plurality of solar panels to a firstpredetermined angle with respect to a first major side of a satellitebody while maintaining a first stacked configuration. In block 1506, themethod 1500 includes releasing, at the first predetermined angle, and bya first cam and a first stop mechanism, a second constrained end of amiddle and outermost panel of the plurality of solar panels. In block1508, the method 1500 includes rotating the released portion of themiddle and outermost panel to a second predetermined angle whilemaintaining a second stacked configuration of the middle panel and theoutermost panel. In block 1510, the method 1500 includes releasing, atthe second predetermined angle, and by a second cam and a second stopmechanism, a third constrained end of the outermost panel.

Certain example implementations can include deploying, from a stowedposition against a first major side of a satellite body, a plurality ofsolar panels having a deployable innermost, middle, and outermost panel,wherein the deployable outermost panel of the plurality of solar panelsis disposed in a first stacked configuration between the deployableinnermost panel and the deployable middle panel during the stowedposition, wherein the plurality of solar panels are configured to berotatably attached to one another therebetween by correspondingfirst-type hinge mechanisms, wherein the deployable innermost panel ofthe plurality of solar panels is configured to attach to the satellitebody at a first edge region by one or more second-type hinge mechanisms,and wherein a hold down and release mechanism (HDRM) disposed at asecond edge region of the satellite body is configured to constrain andselectively release the plurality of solar panels from the stowedposition, the deploying including:

In accordance with an example implementation of the disclosedtechnology, the deployable innermost panel is an outermost panel in thestowed position.

In certain example implementations, the deploying results in a linearsolar array comprising the deployable innermost, middle, and outermostpanel.

In certain example implementations, the axis of the linear array isdisposed about 135 degrees from respective faces of the first major sideand a second major side of the satellite body.

According to an example implementation of the disclosed technology, thefirst-type hinge mechanism is characterized by a rotation angle range ofabout 0 degrees when stowed to about 180 degrees when deployed.

In certain example implementations, the second-type hinge mechanism ischaracterized by a rotation angle range of about 0 when stowed to about135 degrees when deployed.

In certain example implementations, the HDRM is disposed at a fourthedge region of the satellite body, and a first combined rotatable edgeof the plurality of solar panels in the stowed position is disposedopposite the second-type hinge mechanism.

According to an example implementation of the disclosed technology, thesecond-type hinge mechanism constrains the deployed linear array by ahinge spring and a hinge stop.

In certain example implementations, at least one of the plurality ofsolar panels is configured with substantially the same width as acorresponding width of the first major side of the satellite body.

In certain example implementations, at least one panel of the pluralityof solar panels includes solar cells on back and front sides.

According to an example implementation of the disclosed technology, thefirst stop mechanism is configured to engage with the first cam tocontrol a deployment sequence of the plurality of solar panels.

According to an example implementation of the disclosed technology, thesecond cam and a second stop are configured to engage during rotation toprevent the unfolding of the deployable outermost panel until thedeployable middle panel has rotated beyond the second predeterminedangle.

Certain example implementations of the disclosed technology can includeinstalling a fixed solar panel on one or more of a second major side anda fourth major side of the satellite body.

In certain example implementations, the rotating is performed by storedforce of a hinge spring of the one or more of the first-type andsecond-type hinge mechanisms.

FIG. 16 is flow diagram of a method 1600, according to an exampleimplementation of the disclosed technology. In block 1602, the method1600 includes securing one or more panels of a satellite solar arrayagainst a body of the satellite with a burn wire. In block 1604, themethod 1600 includes selectively energizing, with an activation circuit,at least two cylindrical burn resistors in a contact with the burn wire.In block 1606, the method 1600 includes severing the burn wire byselectively energizing the activation circuit. In block 1608, the method1600 includes releasing by the severing, at least the secured end of theone or more panels the satellite solar array. In block 1610, the method1600 includes planar unrolling the satellite solar array into a deployedconfiguration.

Certain example implementations of the disclosed technology can includeretaining, by restraining hooks, the burn wire against the at least twocylindrical burn resistors.

According to an example implementation of the disclosed technology, theselectively energizing directs current through the at least twocylindrical burn resistors to sever the burn wire.

Certain example implementations of the disclosed technology includelaunching the satellite into orbit.

Certain example implementations of the disclosed technology includethermally isolating the two or more cylindrical burn resistors by acircuit board to contain least a portion of heat generated by the atleast two cylindrical burn resistors.

Certain example implementations of the disclosed technology can includesecuring the circuit board to at least one panel of the satellite solararray.

Certain example implementations of the disclosed technology can includemechanically coupling the two cylindrical burn resistors to the circuitboard.

In accordance with an example implementation of the disclosedtechnology, the burn wire may be a monofilament.

Certain implementations of the disclosed technology may include a systemthat includes a satellite body. The satellite body may define: first,second, third, and fourth major sides; with first, second, third andfourth edge regions disposed between the corresponding first and second,second and third, third and fourth, and fourth and first major sides.The system includes a first solar array having a plurality of panelsattached to one another therebetween by corresponding first-type hingemechanisms. A first panel of the plurality of panels is attached to thesatellite body at the first edge region by one or more second-type hingemechanisms, and at least the first panel of the plurality of panels isreleasably constrained by a hold down and release mechanism (HDRM). Thefirst solar array is configured to be foldable (planar rollable) into astowed configuration, and unfoldable (planar unrollable) into a deployedconfiguration.

In certain example implementations, the first-type hinge mechanism ischaracterized by a rotation angle range of about 0 when stowed to about180 degrees when deployed.

In certain example implementations, the second-type hinge mechanism ischaracterized by a rotation angle range of about 0 when stowed to about135 degrees when deployed.

According to an example implementation of the disclosed technology, theHDRM may be configured to constrain the plurality of panels of the firstsolar array against the first major side while in the stowedconfiguration, and release a first combined rotatable edge of the stowedand rolled-up plurality of panels during deployment.

In certain example implementations, the HDRM may be disposed at thefourth edge region, and wherein the first combined rotatable edge of thestowed and rolled-up plurality of panels is disposed opposite thesecond-type hinge mechanism.

Certain example implementations of the disclosed technology may includea second solar array having a second plurality of panels attached to oneanother therebetween by corresponding first-type hinge mechanisms. In anexample implementation, the first panel of the second plurality ofpanels may be attached to the satellite body at the third edge region byone or more second-type hinge mechanisms. In certain exampleimplementations, the second solar array may be configured to be foldableinto a stowed configuration, and unrollable into a deployed linearconfiguration.

Certain example implementations of the disclosed technology may includea second HDRM configured to constrain the second plurality of panels ofthe second solar array against the third major side while in the stowedconfiguration, and release a second combined rotatable edge of thestowed and rolled-up second plurality of panels for deployment.

According to an example implementation of the disclosed technology, thesecond HDRM and the second combined rotatable edge may be disposed atthe second edge region.

According to an example implementation of the disclosed technology, andwhile in the stowed configuration, the first panel of the plurality ofpanels of the first solar array may be exposed to an outer portion ofthe system, a second panel of the plurality of panels of the first solararray may be disposed directly adjacent to the first major side; and athird panel of the plurality of panels of the first solar array may bedisposed between the first panel and the second panel.

According to an example implementation of the disclosed technology, andwhile in the deployed configuration, the plurality of panels may beconfigured as a linear array.

In certain example implementations, an axis of the deployed linear arraymay be disposed about 135 degrees from respective faces of the firstmajor side and the second major side.

In certain example implementations, the second-type hinge mechanism mayconstrain the deployed linear array by a hinge spring and a hinge stop.

According to an example implementation of the disclosed technology, atleast one of the plurality of panels may be configured withsubstantially the same width as a corresponding width of the first majorside.

In certain example implementations, at least one panel of the pluralityof panels may include solar cells on back and front sides.

In certain example implementations, the first plurality of panels caninclude two or more solar cell panels.

Certain example implementations of the disclosed technology can includea stop mechanism adjacent the first edge and configured engage with acam to control a deployment sequence of the plurality of panels.

Certain example implementations of the disclosed technology can includea second stage mechanism including a second cam and a second stopconfigured to engage during rotation to prevent the unfolding of anoutermost panel of the plurality of panels until a middle panel of theplurality of panels has rotated enough beyond a predetermined angle.

According to an example implementation of the disclosed technology, oneor more of the second major side and fourth major side of the satellitebody may be configured with a fixed first solar panel.

In certain example implementations, at least one outermost panel of theplurality of panels is disposed and stacked between an innermost paneland a middle panel of the plurality of panels during stowing.

Certain example implementations of the disclosed technology can includeone or more circuits for selectively actuating the HDRM for deploymentof the plurality of panels.

Certain implementations of the disclosed technology may include hingeassembly configured to rotatingly join an inboard panel with an outboardpanel of a deployable satellite solar array. The hinge assemblyincludes: a first rotor having a pin portion, a rotor endstop portion,and a first rotor body portion. The first rotor body portion isconfigured to attach to the outboard panel. The hinge assembly includesa first bushing having an inner surface configured to rotatingly engagewith an outer surface of the pin portion of the first rotor. The hingeassembly includes a stator having a first bore, a stator endstopportion, a stator body portion, and a second bore. An inner surface ofthe first bore is configured to engage with an outer surface of thefirst bushing, and the stator endstop portion is configured to restricta relative rotation between the inboard and outboard panel when engagedwith the rotor endstop portion. The stator body portion is configured toattach to the inboard panel. The hinge assembly includes a secondbushing having an outer surface configured to engage with an innersurface of the second bore of the stator. The hinge assembly includessecond rotor having a second rotor pin portion, a spring retentionportion and a second rotor body portion. An outer surface of the secondrotor pin portion is configured to rotatingly engage with an innersurface of the second bushing, and the second rotor body portion isconfigured to attach to the outboard panel. The hinge assembly includesa spring having a coiled portion, a first spring end, and a secondspring end. The coiled portion is configured to surround the springretention portion of the second rotor, the first spring end isconfigured to engage with the stator, and the second spring end isconfigured to engage with the second rotor.

According to an example implementation of the disclosed technology, theoutboard panel may be configured to engage with the bushing to act as arotating pin with rotational range constrained by the mechanical endstop.

In certain example implementations, the inboard panel may include a bodyportion of the satellite.

In certain example implementations, the first and second bushings may beshaped as hollow cylinders.

In certain example implementations, the stator can include a firstrecess for constraining the first spring leg. According to an exampleimplementation of the disclosed technology, the first recess mayconstrain the first spring leg between the stator and the inboard panel.

In certain example implementations, the second rotor can include asecond recess for constraining the second spring leg.

In accordance with certain example implementations of the disclosedtechnology, the second recess may constrain the second spring legbetween the second rotor and the outboard panel.

According to an example implementation of the disclosed technology, thesecond bushing may include a flange configured to rotatingly engage witha portion of the spring.

In certain example implementations, the inner surface of the first boremay be configured to rotatingly engage with the outer surface of thefirst bushing.

In certain example implementations, the second bushing outer surface isconfigured to rotatingly engage with the inner surface of the secondbore of the stator.

According to an example implementation of the disclosed technology, thespring may be configured to provide stored energy and actuation force torotate the outboard panel with respect to the inboard panel.

In certain example implementations, spring may be configured to retain arelative rotation of the outboard panel with respect to the inboardpanel in a deployed position when the rotor endstop portion engages withthe stator endstop portion.

According to an example implementation of the disclosed technology, therotor endstop portion may be configured to allow a rotation rangebetween the inboard panel and the outboard panel of about 0 to about 135degrees before engaging with the stator endstop portion.

In certain example implementations, the inboard panel can include aportion of the satellite body.

In accordance with certain example implementations of the disclosedtechnology, the rotor endstop portion may be configured to allow arotation range between the inboard panel and the outboard panel of about0 to about 180 degrees before engaging with the stator endstop portion.

In certain example implementations, the hinge assembly may be configuredto fit within a defined volume along respective edges of the inboard andoutboard panels without intruding onto a volume defined by theassociated panels or adjacent hinges.

In certain example implementations, low-profile screws may be used tofasten the stator to the inboard panel.

In certain example implementations, at least one of the first and secondbushings may be configured to impart rotational resistance to therespective first and second rotors.

Certain implementations of the disclosed technology may include holddown and release mechanism (HDRM) configured to secure and selectivelyrelease one or more panels of a deployable satellite solar array. TheHDRM can include: a circuit board electrically coupled to an activationcircuit; at least two cylindrical burn resistors electrically coupled tothe circuit board; a burn wire (such as monofilament) in contact withthe at least two cylindrical burn resistors; and two or more restraininghooks configured to retain the burn wire against the at least twocylindrical burn resistors. The activation circuit is configured toselectively direct current through the at least two cylindrical burnresistors to sever the burn wire.

In certain example implementations, the circuit board may bemechanically secured to at least one panel of the deployable satellitesolar array. According to an example implementation of the disclosedtechnology, the circuit board may be configured as a thermal isolator toprevent at least a portion of heat generated by the at least twocylindrical burn resistors from spreading to the at least one panel ofthe deployable satellite solar array. In certain example implementationsof the disclosed technology, the circuit board may be mechanicallysecured to at least one panel of the deployable satellite solar array.In certain example implementations, the circuit board may bemechanically secured to a body portion of the satellite.

In certain example implementations, the at least two cylindrical burnresistors may be mechanically coupled to the circuit board.

According to an example implementation of the disclosed technology, therestraining hooks may be mounted to the circuit board.

In certain example implementations, the burn wire may be secured to abody portion of the satellite by at least one fastener. In certainexample implementations, one end of the burn wire may be secured toanother end of the burn wire.

Certain example implementations of the disclosed technology can includea stow detect switch configured to provide indications of the held anddeployed state of the deployable satellite solar array.

In accordance with certain example implementations of the disclosedtechnology, cylindrical burn resistors 1102 may be rated atapproximately 15 ohms. In certain example implementations, the burnresistors 1102 may be selected, as appropriate, to range in resistancefrom about 5 ohms to about 20 ohms, depending on the excitation voltageand current capability applied during deployment. In certain exampleimplementations of the disclosed technology, the burn resistors 1102 maybe cylindrical thin-film MELF resistors. In certain exampleimplementations, the burn resistors 1102 may be rated at 250 mW. Incertain example implementations, the burn resistors 1102 may beselected, as appropriate, from a power rating ranging from about 200 mWto about 500 mW. In certain example implementations, the burn resistors1102 may be in the form of a 0204 package.

In accordance with certain example implementations of the disclosedtechnology, activation circuit is configured to selectively provide 8 ormore volts to one of the at least two cylindrical burn resistors 1102.In certain example implementations, the activation circuit may beconfigured to selectively provide a voltage ranging from approximately 4volts to approximately 12 volts to one of the at least two cylindricalresistors 1102.

In certain example implementations, a hinge spring provides a force todeploy the one or more panels when the burn wire is severed.

In the preceding description, numerous specific details have been setforth. However, it is to be understood that implementations of thedisclosed technology may be practiced without these specific details. Inother instances, well-known methods, structures, and techniques have notbeen shown in detail in order not to obscure an understanding of thisdescription. References to “one implementation,” “an implementation,”“example implementation,” “various implementations,” etc., indicate thatthe implementation(s) of the disclosed technology so described mayinclude a feature, structure, or characteristic, but not everyimplementation necessarily includes the particular feature, structure,or characteristic. Further, repeated use of the phrase “in oneimplementation” does not necessarily refer to the same implementation,although it may.

As used herein, unless otherwise specified the use of the ordinaladjectives “first,” “second,” “third,” etc., to describe a commonobject, merely indicate that different instances of like objects arebeing referred to, and are not intended to imply that the objects sodescribed must be in a sequence, either temporally, spatially, inranking, or in any other manner.

While certain implementations of the disclosed technology have beendescribed regarding what is presently considered to be the mostpractical and various implementations, it is to be understood that thedisclosed technology is not to be limited to the disclosedimplementations, but on the contrary, is intended to cover variousmodifications and equivalent arrangements included within the scope ofthe appended claims. Although specific terms are employed herein, theyare used in a generic and descriptive sense only and not for purposes oflimitation.

This written description uses examples to disclose certainimplementations of the disclosed technology, including the best mode,and to enable any person skilled in the art to practice certainimplementations of the disclosed technology, including making and usingany devices or systems and performing any incorporated methods. Thepatentable scope of certain implementations of the disclosed technologyis defined in the claims, and may include other examples that occur tothose skilled in the art. Such other examples are intended to be withinthe scope of the claims if they have structural elements that do notdiffer from the literal language of the claims, or if they includeequivalent structural elements with insubstantial differences from theliteral language of the claims.

1. A hold down and release mechanism (HDRM) configured to secure andselectively release one or more panels of a deployable satellite solararray, the HDRM comprising: a circuit board electrically coupled to anactivation circuit; at least two cylindrical burn resistors electricallycoupled to the circuit board; a burn wire in contact with the at leasttwo cylindrical burn resistors; two or more restraining hooks configuredto retain the burn wire against the at least two cylindrical burnresistors; wherein the activation circuit is configured to selectivelydirect current through the at least two cylindrical burn resistors tosever the burn wire.
 2. The HDRM of claim 1, wherein the circuit boardis mechanically secured to at least one panel of the deployablesatellite solar array.
 3. The HDRM of claim 2, wherein the circuit boardis configured as a thermal isolator to prevent at least a portion ofheat generated by the at least two cylindrical burn resistors fromspreading to the at least one panel of the deployable satellite solararray.
 4. The HDRM of claim 1, wherein the circuit board is mechanicallysecured to at least one panel of the deployable satellite solar array.5. The HDRM of claim 1, wherein the circuit board is mechanicallysecured to a body portion of the satellite.
 6. The HDRM of claim 1,wherein the least two cylindrical burn resistors are mechanicallycoupled to the circuit board.
 7. The HDRM of claim 1, wherein therestraining hooks are mounted to the circuit board.
 8. The HDRM of claim1, wherein the burn wire is secured to a body portion of the satelliteby at least one fastener.
 9. The HDRM of claim 1, wherein one end of theburn wire is secured to another end of the burn wire.
 10. The HDRM ofclaim 1, further comprising a stow detect switch configured to provideindications of the held and deployed state of the deployable satellitesolar array.
 11. The HDRM of claim 1, wherein the at least twocylindrical resistors are characterized by 15 ohms and rated for 250 mWin a 0204 package.
 12. The HDRM of claim 1, wherein the activationcircuit is configured to selectively provide at least 8 volts to one ofthe at least two cylindrical resistors.
 13. The HDRM of claim 1, whereina hinge spring provides a force to deploy the one or more panels whenthe burn wire is severed.
 14. A method of deploying a satellite solararray, the method comprising: securing one or more panels of thesatellite solar array against a body of the satellite with a burn wire;selectively energizing, with an activation circuit, at least twocylindrical burn resistors in a contact with the burn wire; severing theburn wire by the selectively energizing; releasing by the severing, atleast the secured end of the one or more panels the satellite solararray; and planar unrolling the satellite solar array into a deployedconfiguration.
 15. The method of claim 14, further comprising retainingby restraining hooks, the burn wire against the at least two cylindricalburn resistors.
 16. The method of claim 14, wherein the selectivelyenergizing directs current through the at least two cylindrical burnresistors to sever the burn wire.
 17. The method of claim 14, furthercomprising launching the satellite into orbit.
 18. The method of claim14, further comprising thermally isolating the two or more cylindricalburn resistors by a circuit board to contain least a portion of heatgenerated by the at least two cylindrical burn resistors.
 19. The methodof claim 18, further comprising securing the circuit board to at leastone panel of the satellite solar array.
 20. The method of claim 1,further comprising mechanically coupling the two cylindrical burnresistors to the circuit board.